AIRCRAFT CAPABLE OF HOVERING AND RELATIVE CONTROL METHOD

Description

CROSS-REFERENCE T0 RELATED APPLICATIONS

This Patent Application claims priority from European Patent Application No. 22188969.4 filed on Aug. 5, 2022, the entire disclosure of which is incorporated herein by reference.

TECHNICAL FIELD

The present invention relates to an aircraft capable of hovering, for example a helicopter or a convertiplane or a heliplane.

The present invention also relates to a method for controlling an aircraft capable of hovering.

BACKGROUND

In the aviation sector, aeroplanes are normally used for high cruising speeds, in particular greater than 150 knots and high altitudes, e.g. above 30,000 feet. At high cruising speeds and altitudes, aeroplanes use fixed wings to generate the lift necessary to sustain the plane itself. A sufficient value of this lift can only be obtained by accelerating the aeroplane on runways of considerable length. These runways are also necessary to allow the same aeroplanes to land.

In contrast, the helicopters normally have lower cruising speeds than those of the aeroplanes and generate the necessary lift for sustenance through the rotation of the blades of the main rotor. As a result, helicopters can land/take off without the need for a horizontal speed and using particularly small surfaces. Moreover, helicopters are able to hover and to fly at relatively low altitudes and speeds, resulting thus as particularly manoeuvrable and suitable for demanding manoeuvres such as rescuing people in the mountains or at sea.

Nevertheless, helicopters have inherent limitations in terms of maximum operating altitude, which is around 20000 feet, and of maximum operating speed, which cannot exceed 150 knots.

In order to meet the demand for aircrafts capable of presenting the same manoeuvrability and comfort of use as the helicopter and at the same time overcoming the inherent limitations mentioned above, convertiplanes and heliplanes are known.

An example of a convertiplane is described in patent application U.S. Pat. No 10,011,349.

In greater detail, the convertiplane described in the aforesaid application essentially comprises:

• • a fuselage extending along a first longitudinal axis; and
  • a wing protruding cantilevered and formed by a pair of half-wings arranged on respective mutually opposite parts of the fuselage, and having respective free ends opposite to the fuselage and aligned along a second transverse axis that is substantially orthogonal to the first longitudinal axis.

The convertiplane further comprises:

• • a pair of nacelles housing the respective motors; and
  • a pair of rotors that are rotatable around respective third axes and operatively connected to respective motors.

The rotors are tiltable with respect to the wing around a fourth axis, preferably parallel to the second axis.

The convertiplanes are also able to selectively assume:

• • a first “aeroplane” configuration, in which the rotors are arranged with respective third axes that are substantially parallel to the first axis of the same convertiplane and coaxial to the respective engines; or
  • a second “helicopter” configuration, in which the rotors are arranged with the respective third axes that are substantially vertical and orthogonal to the first axis of the convertiplane and orthogonal to the respective motors.

Thanks to the possibility of tilting the rotors, the convertiplanes are able to take off and land like a helicopter, i.e. in a direction substantially perpendicular to the first longitudinal axis of the convertiplane, without the need for a runway.

Furthermore, the convertiplanes are able to take off and land on rough terrains and without generating a noise level incompatible with an urban settlement.

In addition, the convertiplanes are capable of hovering when arranged in the helicopter configuration.

Furthermore, the convertiplanes can reach and maintain cruising speeds of approximately 250-300 knots and flight altitudes of the order of 30000 feet when arranged in the airplane configuration.

This cruising speed is well above the value of about 150 knots that defines the maximum cruising speed of the helicopters.

Similarly, the above altitude is well above the one typical of the helicopters and allows convertiplanes arranged in an airplane configuration to avoid the clouds and atmospheric disturbances characteristic of lower altitudes.

The heliplanes, such as, for example, the EUROCOPTER X-3 aircraft comprise, in addition to the components commonly found in a known helicopter such as a main rotor with vertical axis, a pair of half-wings protruding cantilevered from respective parts of the fuselage of the heliplane along a fourth transverse axis substantially orthogonal to a fifth longitudinal axis of the aircraft and to the axis of rotation of the main rotor.

In more detail, each of the half-wings carries a respective propeller which comprises, in a known manner, a drive shaft operable by a relative motor and a plurality of blades articulated on the drive shaft itself.

In particular, each drive shaft is rotatable around a relative sixth axis substantially parallel to the longitudinal axis of the heliplane, i.e., a horizontal axis.

The heliplane is therefore able, in the same way as the convertiplane, to take off and land in a vertical direction by means of the main rotor and to fly in forward flight by means of the propellers and the aforesaid half-wings.

During the forward flight, the main rotor rotates idly while the thrust is generated by the propellers.

Electrically-propelled or hybrid-propelled aircrafts are known wherein at least one propulsion element (e.g. a propeller or rotor) is operable by battery-powered electric motors.

In such aircrafts, the temperature of the batteries must be strictly maintained within a temperature range. In fact, an uncontrolled increase in the temperature of the batteries could lead to a condition known as “thermal runaway”, in which flames are formed or explosions are triggered, and which can have disastrous consequences for the entire aircraft.

In the field of the electrically-or hybrid-propelled aeroplanes, a number of battery cooling solutions have been developed, including those shown in U.S. Pat. No. 10,177,424, WO-A-2021/064374, EP-B-3176851, U.S. Pat. No. 9,415,878 and U.S. Pat. No. 3,957,230.

According to these solutions, a battery of the aeroplane is arranged inside a cooling duct obtained in the fuselage or at a wing and is impinged by an air flow due to the motion of the aircraft itself.

However, such solutions developed specifically for aeroplanes do not make it possible to effectively regulate the temperature of the batteries of the aircraft capable of hovering.

WO-A1-2021222528 discloses an aircraft nacelle having a first and second heat exchanger section to cool aircraft during different modes. Additionally, a fan and other components are configured to maximize efficiency and cooling capacity during a plurality of operating conditions.

CN-A-113193209 discloses a fixed-wing unmanned aerial vehicle air cooling type fuel cell double-stack integrated power system. The system comprises a fuselage, a power motor, a high-pressure hydrogen storage tank, wings, two air cooling type fuel cell stacks symmetrically arranged in the middle of the fuselage or on the wings, and two heat dissipation systems corresponding to the air cooling type fuel cell stacks; an air flow channel of the electric pile is a parallel wave-shaped flow channel; when the galvanic pile is arranged in the middle of the machine body, the cooling system further comprises a cooling fan; when the galvanic pile is arranged on the wings, the power system is suitable for the unmanned aerial vehicle with auxiliary propellers on the wings, and the heat dissipation system further comprises a wing front air guide cover; when the unmanned aerial vehicle is started or flies at a low speed, a medium speed and a high speed, the reaction temperature of the galvanic pile is controlled to be in an ideal temperature interval through different heat dissipation modes.

WO-A1-2021106549 discloses a multi-rotor helicopter having a fuselage and a plurality of fan units. Each of the fan units is equipped with a circular fan frame, a rotating blade, and a drive-system cooling unit. Each drive-system cooling unit has: an accommodation container that accommodates at least one of a drive unit, a driver, and a power source; a cooling fan that supplies cooling air to the accommodation container; an intake flow path that guides air from the cooling fan toward the accommodation container; and an exhaust flow path that discharges air that has passed through the accommodation container. The discharge flow path discharges air that has passed through the accommodation container in the tangent direction of the fan frame.

In aircrafts capable of hovering, in fact, the risk that the temperature of the batteries increases in an uncontrolled way is particularly high, especially while hovering. Specifically, in this flying condition, the air flow rate that invests the aircraft and that would be destined for heat exchange with the batteries is much lower than during the forward flight.

There is a perceived need in the industry to realize an aircraft capable of hovering, wherein the temperature of the batteries can be regulated efficiently.

Aim of the present invention is to realize an aircraft capable of hovering, which allows to meet the need specified above in a simple and economical way.

SUMMARY

According to the invention, this aim is achieved by an aircraft capable of hovering as claimed in Claims 1 and 7 and by a method for controlling an aircraft capable of hovering as claimed in Claims 14 and 16.

BRIEF DESCRIPTION OF THE DRAWINGS

For a better understanding of the present invention, a preferred non-limiting embodiment is described below, purely by way of example and with the aid of the attached drawings, wherein:

FIG. 1A is a side view of an aircraft according to the present invention in a first operating configuration, comprising a cooling system and with parts removed for clarity's sake;

FIG. 1B is a side view of the aircraft shown in FIG. 1A in a second operating configuration;

FIG. 2 is a perspective view of the aircraft of FIGS. 1A and 1B in an enlarged scale and with parts removed for clarity's sake;

FIG. 3 is an exploded view of the cooling system of FIGS. 1A and 1B with parts removed for clarity's sake;

FIG. 4 is a top view of the cooling system of FIGS. 1A, 1B and 3 with parts removed for clarity's sake;

FIG. 5 is a detail of FIG. 4 on a greatly enlarged scale and with a partial section; and

FIG. 6 is a section along line VI-VI in FIG. 1A with parts removed for clarity's sake.

DESCRIPTION OF EMBODIMENTS

With reference to FIGS. 1A and 1B, 1 denotes an aircraft capable of hovering with at least partly electric propulsion.

In greater detail, the aircraft 1 is a convertiplane selectively switchable between:

• • a first configuration (FIG. 1A), in which it is in a forward flight condition and proceeds along a predominantly horizontal trajectory; and
  • a second configuration (FIG. 1B), in which it performs a hovering manoeuvre or moves forward along a predominantly vertical trajectory.

It must be specified that in the following present disclosure, expressions such as “upper”, “lower”, “at the front”, “at the back” and the like are used with reference to normal forward flight or “hovering” conditions of the aircraft 1.

It is possible to identify a triplet of axes integral to the aircraft 1 and originating at a centre of gravity of the aircraft 1 itself formed by:

• • a longitudinal axis Y of the same aircraft 1;
  • an axis X orthogonal to the axis Y; and
  • an axis Z orthogonal to the axes X, Y.

It is also possible to define a median plane M of the aircraft 1 with respect to the axis X and directed parallel to the axis Y.

The aircraft 1 essentially comprises:

• • a fuselage 2 elongated along the longitudinal Y axis;
  • a plurality of rotors 3a, 3b, 4 that are rotatable around respective rotation axes B, C, D, E, F, G with respect to the fuselage 2;
  • electrical drive means—not shown—adapted to rotate at least one of the rotors 3a, 3b and 4;
  • a plurality of batteries 9 adapted to electrically power the electrical drive means; and
  • a cooling system 10 of the batteries 9.

In detail, the cooling system 10 is adapted to regulate the temperature T of the batteries 9 by means of the heat exchange between an air flow taken from the outside and the batteries 9.

The aircraft 1 could further comprise one or more thermal motors for driving one or more of the rotors 3a, 3b and 4. In other words, the aircraft 1 could be with hybrid propulsion.

As shown in FIGS. 1A and 1B, the fuselage 2 defines a nose 5 and a tail 6 of the aircraft 1, which are opposite to each other along the longitudinal axis Y. In addition, the fuselage 2 comprises a belly 7, which is interposed between the nose 5 and the tail 6 along the longitudinal axis Y.

In detail, the belly 7 is adapted to be facing towards the ground during the normal operation of the aircraft 1.

With reference to the normal forward flight operating conditions, the aircraft 1 proceeds in a direction oriented from the tail 6 to the nose 5 with a forward speed v with respect to the ground (FIG. 1A).

In greater detail, the aircraft 1 comprises:

• • a pair of half-wings 8 extending cantilevered from respective mutually opposite sidewalls 62 of the fuselage 2 and transversely to the axis Y (FIG. 2);
  • a pair of rotors 3a that are rotatable about respective fixed axes B, C with respect to the fuselage 2;
  • a pair of rotors 3b that are rotatable about respective fixed axes D, E with respect to the fuselage 2; and
  • a pair of rotors 4 that are rotatable about respective axes F, G and tiltable with respect to an axis H between a first position assumed when the aircraft 1 is in the first configuration and a second position assumed when the aircraft 1 is in the second configuration.

In detail, the axis H is parallel to the axis X.

The axes B, C and the axes D, E lie on two respective planes parallel to the axes X and Z.

In addition, the axes B and C are incident with each other and are tilted with respect to the axis Z, in particular at a point arranged above the belly 7. In greater detail, the axes B and C are both tilted by 10° with respect to the axis Z.

Similarly to the axes B and C, the axes D and E are incident with each other and tilted with respect to the axis Z, in particular at a point arranged above the belly 7. In greater detail, the axes D and E are both tilted by 10° with respect to the axis Z.

The rotors of each pair of rotors 3a and 3b are arranged symmetrically with respect to the median plane M. In addition, the pair of rotors 3a is arranged at the nose 5, the pair of rotors 3b is arranged at the tail 6, and the pair of rotors 4 is interposed between the pair of rotors 3a and the pair of rotors 3b along the longitudinal axis Y.

The axes F, G are arranged orthogonally to the axes B, C; D, E and parallel to the axis Y when the rotors 4 are arranged in the first position.

The axes F and G are arranged parallel to the axis Z when the rotors 4 are arranged in the second position (FIG. 1B).

Preferably, the rotors 3a, 3b and 4 are with fixed pitch.

In the embodiment shown, each of the rotors 3a, 3b and 4 is driven by a respective electric motor of the electric drive means. In detail, each electric motor is operable independently of the other electric motors.

The aircraft 1 further comprises a control unit 60 (only schematically shown in FIG. 5) receiving as input a plurality of control signals provided by the crew, by an autopilot or a remote control system, and programmed to provide as output a plurality of commands to command the rotors 3a, 3b and 4 so that they provide desired values of the relative thrusts. In greater detail, the control unit 60 is programmed to command the rotors 3a, 3b and 4 to generate respective thrusts independent of each other.

Referring to FIG. 2, the cooling system 10 comprises:

• • an opening 20 for the air to enter;
  • a plurality of openings 21 for the air to escape; and
  • a passage 22, which extends through the fuselage 2 and fluidly connects the opening 20 with the openings 21.

In particular, the batteries 9 are placed within the passage 22 and are fluidically interposed between the opening 20 and at least part of the openings 21.

Advantageously, the cooling system 10 comprises two fans 23 adapted to increase the kinetic energy of the air contained in the passage 22 (FIG. 5); these fans 23 are adapted to be operated when the forward speed v of the aircraft 1 with respect to the ground is lower than a speed threshold value v0 and/or when the temperature T of the batteries 9 exceeds a temperature threshold value T0.

For example, the temperature threshold value T0 is lower than 75° C. Preferably, the temperature threshold value T0 is equal to 60° C.

The control unit 60 is also operatively connected to the fans 23 to control their operation, i.e. to command the rotation of the fans 23 around respective rotation axes I, J (FIG. 5).

In detail, when the forward speed v is greater than the speed threshold value v0 (e.g., during the forward flight) and/or the temperature T of the batteries is lower than the temperature threshold value T0, the control unit 60 is adapted to deactivate or keep the fans 23 deactivated. In this condition, the batteries 9 are cooled by the flow of air entering in the cooling system 10 through the opening 20 due to the effect of the relative motion of the aircraft 1 with respect to the air in which it is immersed. This type of cooling is called “ram ventilation”.

Conversely, when the forward speed v is lower than the speed threshold value v0 (e.g., while hovering) and/or the temperature T of the batteries is greater than the temperature threshold value T0, the control unit 60 is adapted to activate the fans 23. In this condition, the batteries 9 are cooled by the flow of air entering the cooling system 10 through the opening 20, which is forced by the action of the fans 23.

As shown in FIGS. 1A and 1B, the opening 20 is arranged at the nose 5 and the openings 21 are arranged at the belly 7.

In detail, the opening 20 is centred with respect to the median plane M (FIG. 2).

Preferably, furthermore, the opening 20 comprises:

• • a curved section 20a having an upwardly facing curvature, i.e. towards a portion of the aircraft 1 opposite to the belly 7 along the axis Z;
  • a curved section 20b spaced from the curved section 20a parallel to the axis Z towards the belly 7 and also having an upwardly facing curvature; and
  • two curved sections 20c, 20d, in particular in the form of an arc of circumference, which connect the curved sections 20a, 20b at their respective opposite ends parallel to the axis X.

In other words, the opening 20 has a curved elliptical shape, i.e. a bean shape.

The aircraft 1 further comprises three containers 41, 42, 43 defining respective inner volumes 50, within which respective pluralities batteries 9 are contained. The batteries 9 within each inner volume 50 define a plurality of interstices 45 with one another and the respective container 41, 42, 43.

In the embodiment shown, the containers 41, 42 and 43 are parallelepiped-shaped (FIG. 2). In detail, the containers 41, 42 and 43 have a square or substantially square base in a plane parallel to the axes X and Y. In addition, the extension of the containers 41, 42, and 43 parallel to the axis Z is smaller (e.g., ⅕ or ⅙) than the extension of the containers 41, 42, and 43 parallel to the axes X and Y.

The containers 41, 42, 43, moreover, are aligned with each other parallel to the axis Y and are centred with respect to the median plane M.

Preferably, the containers 41, 42 and 43 are identical to each other.

The batteries 9 are shaped like an elongated parallelepiped along a direction K. The batteries 9 are furthermore parallel to each other, i.e. arranged so that the relative directions K coincide, and aligned with each other parallel to the axis X. In the embodiment shown, the directions K are parallel to the longitudinal axis Y. In addition, inside each container 41, 42, 43 the batteries 9 are fixed to each other.

In the embodiment shown, each container 41, 42, 43 contains five batteries 9.

In detail, temperature T of the batteries 9 refers to the temperature at the outer surface of the batteries 9, or in the vicinity of the batteries 9, for example within the containers 41, 42, 43.

As shown in FIGS. 2, 3 and 4, the passage 22 comprises:

• • a duct 30, which extends starting from the opening 20;
  • three ducts 31, 32, 33, which branch off from the duct 30 and which each fluidically connect the duct 30 to the inner volume 50 of a respective container 41, 42, 43; and
  • the interstices 45.

In greater detail, proceeding from the opening 20 along the longitudinal axis Y towards the tail 6, the duct 30 comprises a first section 30a and a second section 30b joined together.

The second section 30b is directed parallel to the longitudinal axis Y and the first section 30a extends obliquely with respect to the second section 30b. In detail, the opening 20 is arranged below the second section 30b with respect to the axis Z.

The second section 30b also has a circular cross-section and the first section 30a has a progressively variable shaped section. In detail, the shape of the cross-section of the first section 30a initially corresponds to the shape of the opening 20 and then connects to the circular section of the second section 30b (FIGS. 4 and 5).

Preferably, moreover, the passage section of the duct 30 has progressively decreasing extension proceeding from the opening 20 along the longitudinal axis Y towards the tail 6.

In greater detail, the duct 31 fluidically connects the second section 30b to the inner volume 50 of the container 41 and is directed substantially parallel to the axis Z. The duct 32 fluidically connects the second section 30b to the inner volume 50 of the container 42. The ducts 31 and 32, moreover, are centred with respect to the median plane M.

The duct 33 fluidically connects the second section 30b to the inner volume 50 of the container 43 and comprises two branches 33a, 33b, which are arranged symmetrically with respect to the median plane M.

The cross-section of the duct 31 has a constant or substantially constant extension parallel to the axis Z. In addition, the ducts 32 and 33 have constant or substantially constant extension along the longitudinal axis Y.

The passage 22 also comprises two auxiliary ducts 34, 35, at which a respective fan 23 is housed (FIG. 5).

Each of said auxiliary ducts 34, 35 comprises respective mutually opposite ends 34a, 34b; 35a, 35b. These ends 34a, 34b; 35a, 35b are directly facing the duct 30 and in fluidic communication therewith (FIGS. 4 and 5).

In greater detail, the auxiliary ducts 34, 35 are directly connected to the second section 30b of the duct 30.

Considering a cross-section of the passage 22 passing through a plane orthogonal to the axis Z, the auxiliary ducts 34 and 35 are U-shaped and are arranged symmetrically with respect to each other with respect to the median plane M (FIG. 5).

Each auxiliary duct 34, 35 has a cross-section having an extension lower than the minimum extension of the cross-section of the duct 30. In addition, the sum of the maximum extensions of the cross-sections of the auxiliary ducts 34 and 35 is lower than the minimum extension of the cross-section of the duct 30.

As shown in FIG. 3, the openings 21 have a rectangular section in a plane orthogonal to the axis Z, are arranged parallel to each other and to the longitudinal axis Y, and are spaced from each other parallel to the axis X.

In the embodiment shown in FIG. 3, the containers 41, 42 and 43 each comprise:

• • a cover 46;
  • a base plate 47; and
  • a set of side walls 48 extending between the cover 46 and the base plate 47 parallel to the axis Z.

The cover 46, the base plate 47 and the set of side walls 48 of each container 41, 42 and 43 define the inner volume 50 of the relative container.

In particular, the openings 21 are obtained at the base plate 47.

Preferably, the covers 46 are fixed to each other and to the ducts 31, 32 and 33 (FIG. 3).

The aircraft 1 also comprises (FIG. 5):

• • sensor means 65 adapted to detect the temperature T of the batteries 9 and operatively connected to the control unit 60;
  • sensor means 70 adapted to detect the forward speed v of the aircraft 1 and operatively connected to the control unit 60.

Preferably, the sensor means 70 comprise a flow meter adapted to detect the flow rate that invests, in use, the aircraft 1 in parallel to a horizontal or substantially horizontal forward direction of the aircraft 1.

The cooling system 10 further comprises means for varying the flow rate of air entering through the opening 20, not shown.

Such flow rate variation means comprise, for example, a valve adapted to partialise the flow rate of entering air and operatively connected to the control unit 60.

In detail, the control unit 60 is programmed to command the partialisation of the flow rate of air entering through the valve when the temperature T of the batteries 9 is lower than a minimum temperature threshold value Tmin, which is lower than the temperature threshold value T0. For example, the minimum temperature threshold value Tmin is equal to 0° C.

The operation of the aircraft 1 according to the invention is described below.

In use, the aircraft 1 lands and takes off arranged in the second configuration with the rotors 4 arranged in the second position (FIG. 1B). In this second configuration, the lift required to sustain the aircraft 1 is provided by the rotors 3a, 3b and 4.

During the transition from the first to the second configuration of the aircraft, the control unit 60 is programmed to reduce the thrusts generated by the rotors 3a and 3b as the axes F, G of the rotors 4 progressively approach a condition of parallelism with the axis Y and the speed v of the aircraft 1 increases.

The aircraft 1 moves forward at cruising speed in the first configuration with the rotors 4 arranged in the first position (FIG. 1A). In this first configuration, the lift required to sustain the aircraft 1 is provided for the most part at least by the half-wings 8 and/or by other aerodynamic surfaces arranged along the aircraft 1. The rotors 3a and 3b can be deactivated if necessary.

During use, the sensor means 65 detect the temperature T of the batteries 9 and/or the sensor means 70 detect the forward speed v.

If the forward speed v is greater than the speed threshold value v0 and/or the temperature T is less than the temperature threshold value T0, the control unit 60 deactivates the fans 23 or keeps them deactivated.

In detail, when the fans 23 are deactivated, the air enters the cooling system 10 through the opening 20 due to the effect of the motion of the aircraft 1, crosses the duct 30 and is distributed among the ducts 31, 32 and 33 reaching the containers 41, 42 and 43. Within the inner volumes 50 of the containers 41, 42 and 43, the air flows in the interstices 45, absorbing the heat of the batteries 9, and then escapes from the openings 21.

In greater detail, during the crossing of the passage 22, the air flow transits largely through the second section 30b and minimally through the auxiliary ducts 34 and 35, by virtue of the cross-sectional dimensions of these auxiliary ducts 34, 35 with respect to the cross-sectional dimensions of the second section 30b.

Conversely, if the forward speed v is lower than the speed threshold value v0 (e.g. when aircraft 1 is hovering), or the temperature T exceeds the temperature threshold value T0, the control unit 60 activates the fans 23.

In detail, when the fans 23 are active, the air passes through in order the same ducts it passes through when the fans 23 are deactivated. However, since the fans 23 are activated, the kinetic energy of the air is increased and the forced ventilation of the batteries 9 is achieved.

If during operation of the aircraft 1 the temperature T of the batteries 9 falls below the minimum temperature threshold value Tmin, the control unit 60 commands the partialisation of the entering air flow rate. In this way, the amount of heat removed from the batteries 9 is reduced.

An examination of the characteristics of the aircraft 1 shows the advantages that it allows obtaining.

Since the cooling system 10 comprises the fans 23, which perform the forced ventilation when the forward speed v is lower than the speed threshold value v0 and/or when the temperature T exceeds the temperature threshold value T0, it is possible to effectively regulate the temperature of the batteries 9 of the aircraft 1. This is particularly true when the aircraft 1 is hovering and the flow rate of air entering through the opening 20 is therefore limited or in any case characterized by low kinetic energy.

Since the opening 20 is arranged at the nose 5, it is possible to maximize the flow rate of air entering through the opening 20 itself. At the same time, since the openings 21 are arranged at the belly 7, the flow of air exiting the cooling system 10 does not disturb the aerodynamics of the aircraft 1.

Since the fans 23 are respectively arranged in the auxiliary ducts 34, 35, the fans 23 when they are deactivated do not constitute an obstacle to the transit of air, which passes substantially undisturbed through the duct 30.

It is clear that the aircraft 1 described and shown herein may be subject to modifications and variations without thereby departing from the scope of protection defined by the Claims.

The aircraft 1 could be a helicopter or a helicoplane.

At least some or all of the rotors 3a, 3b and 4 could be with variable pitch.

The passage 22 could comprise a single auxiliary duct 34, 35, or more than two auxiliary ducts 34, 35.

The cooling system 10 could comprise a single fan 23, or more than one fan 23. In particular, the cooling system 10 could comprise more than one fan 23 for each of the auxiliary ducts 34, 35.

The aircraft 1 could comprise one, or two containers 41, 42, 43, or even more than three containers 41, 42, 43. In addition, the containers 41, 42, 43 could not be aligned with each other.

The passage 22 could comprise a single duct 30 fluidly connecting opening 20 to a single container 41, 42 or 43. Preferably, single duct 30 directly fluidly connects opening 20 to the single container 41, 42, 43. In other words, passage 22 could not comprise ducts 31, 32, 33. According to this embodiment, the one or more auxiliary ducts 34, 35 comprise respective first and second ends 34a, 35a, 34b, 35b both directly facing single duct 30.

The directions K of the batteries 9 could be arranged parallel to the axis X and the batteries 9 could be aligned with each other along the longitudinal axis Y. Additionally or alternatively, the openings 21 could be arranged parallel to each other and to the axis X and be spaced apart from each other parallel to the longitudinal axis Y.

The sensor means 65, 70 could be connected directly to the avionic devices on board the aircraft 1 or could be connected to control units other than the control unit 60.